(1) FAA Technical Report ADS-4, Engineering Summary of Airframe Icing Technical Data, December 1963.
(2) Advisory Circular 20-73, Aircraft Ice Protection, April 21, 1971.
(3) Advisory Circular 91-51, Airplane De-ice and Anti-ice Systems, September 15, 1977.
(4) FAA Report RD-77-76, Engineering Summary of Powerplant Icing Technical Data, July 1977.
(5) United States Army Aviation Engineering Flight Activity Reports:
(i) Natural Icing Tests, UH-1H Helicopter, Final Report, June 1974, USAASTA Project No. 74-31.
(ii) Artificial Icing Tests, UH-1H Helicopter, Part 1, Final Report, January 1974, USAASTA Project No. 73-04-4.
(iii) Artificial Icing Tests, UH-1H Helicopter, Part II, Heated Glass Windshield, Final Report, USAASTA Project No. 73-04-4.
(iv) Artificial Icing Tests, Lockheed Advanced Ice Protection System Installed on a UH-1H Helicopter, Final Report, June 1975, USAAEFA Project No. 74-13.
(v) Artificial and Natural Icing Tests for Qualification of the UH-1H, Kit A Aircraft, Letter Report, USAAEFA Project No. 78-21-1.
(vi) Microphysical Properties of Artificial and Natural Clouds and Their Effects on UH-1H Helicopter Icing, Report USAAEFA Project No. 78-21-2.
(vii) Helicopter Icing Spray System (HISS) Nozzle Improvement Evaluation, Final Report, September 1981, USAAEFA Project No. 79-002-2.
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(viii) Artificial and Natural Icing Tests of the YCH-4TD, Final Report, May 1981, USAAEFA Project No. 79-07.
(ix) Limited Artificial Icing Tests of the OV-ID, Letter Report, July 1981, USAAEFA Project No. 80-16, and (Limited Distribution).
(x) JUH-IH Ice Phobic Coating Tests, Final Report, July 1980, USAAEFA Project No. 79-02.
(xi) Artificial and Natural Icing Tests, Production UH-60A Helicopter, Final Report, June 1980, USAAEFA Project No. 79-19.
(xii) Helicopter Icing Spray System (HISS) Evaluation and Improvements, Letter Report, June 1981, USAAEFA Project No. 80-04.
(xiii) Artificial Icing Test of CH-47C Helicopter with Fiberglass Rotor Blades, Final Report, July 1979, USAAEFA Project No. 78-18.
(xiv) Limited Artificial and Natural Icing Tests, Production UH-60A Helicopter (Reevaluation), Final Report, August 1981, USAAEFA Project No. 80-14.
(6) Further Icing Experiments on an Unheated Nonrotating Cylinder, National Research Council, Canada Report LTR-LT-105, dated November 1979, by J.R.
Stallabrass and P.F. Hearty.
(7) Ludlam, F.H., Heat Economy of a Rimed Cylinder, Quarterly Journal, Royal Meteorological Society, Vol. 77, 1951.
(8) U.S. Army AMRDL Reports:
(i) USAAMRDL TR 73-38, Ice Protection Investigation For Advanced Rotary Wing Aircraft, J.B. Werner, August 1973, AD 7711182.
(ii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 1, Design Criteria and Technology Considerations, USAAMRDL - TR-75-34A, Eustis Directorate, U.S. Army Air Mobility R&D Laboratory, November 1975, AD A019044.
(iii) Werner, J.B., The Development of an Advanced Anti-Icing/Deicing Capability for U.S. Army Helicopters, Volume 2, Ice Protection System Application to the UH-1H Helicopter, USAAMRDL - TR-75-34B, Eustis Directorate, U.S. Army Air Mobility Research and Development Laboratory, November 1975, AD A019049.
(iv) USAAMRDL-TR-76-32, Ottawa Spray Rig Tests of an Ice Protection System Applied to the UH-1H Helicopter, November 1976, AD A0034458.
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(v) USARTL-TR-78-48, Icing Tests of a UH-1H Helicopter with an Electrothermal Ice Protection System Under Simulated and Natural Icing Conditions, April 1979.
(vi) USAAMRDL-TR77-36, Final Report, Natural Icing Flights and Additional Simulated Icing Tests of a UH-IH Helicopter Incorporating an Electrothermal Ice Protection System, July 1978, AD A059704.
(9) Technical Feasibility Test of Ice Phobic Coatings for Rain Erosion in Simulated Flight Conditions, U.S. Army Test and Evaluation Command, Final Report, 4-AI-192-IPS-001, August 1980.
(10) Technical Feasibility Test of Ice Phobic Coatings in Simulated Icing Flight Conditions, U.S. Army TECOM, Final Report, 4-CO-160-000-048, September 1980.
(11) Aircraft Icing, NASA Conference Publication 2086, FAA-RD-78-109, July 1978.
(12) Helicopter Icing Review, FAA Technical Center, Final Report, FAA-CT-80-210, September 1980.
(13) National Icing Facilities Requirements Investigation, Final Report, FAA Technical Center, FAA-CT-81-35, March 1981.
(14) Aircraft Icing, AGARD Advisory Report No. 127, November 1978.
(15) Rotorcraft Icing - Review and Prospects, AGARD Advisory Report, AR-166, September 1981.
(16) Advisory Circular 20-117, Hazards Following Ground Deicing and Ground Operations in Conditions Conducive to Aircraft Icing, December 17, 1982.
(17) Olson, W., Experimental Comparison of Icing Cloud Instruments, January 1983, NASA TM 83340.
(18) JUH-1H Redesigned Pneumatic Boot Deicing System Flight Test Evaluation. Hayworth, L., Graham, M., August 1987. USAAEFA Edwards AFB, California. Project No. 83- 13.
(19) An Appraisal of the Single Rotating Cylinder Method of Liquid Water Content Measurement, National Research Council Canada Report LTR-LT-92, dated November 1978, by J.R. Stallabrass.
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40 - 30 - AMBIENT TEMPERATURE - DEG F 20 - 10 -10 - 2000 4000 6000 8000 10,000
PRESSURE ALTITUDE - FT FIGURE AC 29.1419-1 CONTINUOUS ICING - TEMPERATURE VS ALTITUDE LIMITS Figures AC 29.1419 - 1 through 4 represent the approach to a 10,000-foot altitude limit.
See Paragraph b(5)(Ili) for a discussion of this approach.
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40.
30 AMBIENT TEMPERATURE - DEG F 20 • 10 10 2000 4000 6000 8000 10,000
PRESSURE ALTITUDE - FT FIGURE AC 29.1419-2 INTERMITTENT ICING - TEMPERATURE VS ALTITUDE LIMITS Figures AC 29.1419-1 through 4 represent the approach to a 10,000-foot altitude limit.
See Paragraph b(5)(iii) for a discussion on this approach.
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1.0 - LIQUID WATER CONTENT - GRAMS/CUBIC METER 0.8 - SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE
0.6 AMBIENT TEMP °F 15
0.4 *70
0.2 10 20 30 40 50
MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE AC 29.1419-3 CONTINUOUS ICING - LIQUID WATER CONTENT VS. DROP DIAMETER Figures AC 29.1419-1 through 4 represent one approach to a 10,000-foot altitude limit.
See Paragraph b(5) (ili) for a discussion on this approach.
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2.0 SEA LEVEL TO 10,000 FT PRESSURE ALTITUDE LIQUID WATER CONTENT - GRAMS/CUBIC METER 1.6
AMBIENT TEMP
- 32 °F
1.2
- 14 °F
0.8
- 0 °F
0.4 10 20 30 40 50
MEAN EFFECTIVE DROP DIAMETER - MICRONS FIGURE AC 29. 1419-4 INTERMITTENT ICING - LIQUID WATER
CONTENT VS DROP DIAMETER Figures AC 29.1419-1 through 4 represent one approach to a 10,000-foot altitude limit.
See Paragraph b(5) (ili) for a discussion of this approach.
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SUBPART F - EQUIPMENT
MISCELLANEOUS EQUIPMENT AC 29.1431. § 29.1431 ELECTRONIC EQUIPMENT.
a. Background. Th is section contains some specific requirements for electronic equipment in the rotorcraft. The principal requirements of this section are that radio and navigation equipment must be free from hazards, both in themselves and in their effect on any other items installed in the rotorcraft, and that operation of the radio and navigation equipment does not interfere with operation of any other required avionics.
The increased use of complex equipment that integrates communication and navigation functions increases the likelihood of common mode failures resulting in simultaneous loss of communication and navigation functions. Total non- restorable loss of communication and navigation information is considered to be a catastrophic failure condition for IFR operations.
b. Procedures. In showing compliance with this section, tests and analysis should be performed as necessary to determine that:
(1) All radio and navigation equipment is installed and operated in such a manner that it does not result in hazards to the rotorcraft. It also should not have an effect on any other components of the rotorcraft to the extent that it creates a hazardous condition. Consideration should be given to the effects of critical environmental conditions. The environment can easily be the cause of common mode failures.
Temperature extremes in the rotorcraft may exceed the temperature to which the system was qualified. Additional considerations include:
(i) An analysis, per SAE ARP 4761, to assure there is no single condition or fault which can cause multiple channels, systems, circuits, etc. to fail simultaneously.
An example of this could be a common power supply for both communication and navigation functions.
(ii) Addressing each potential common cause fault case and identifying the corresponding mitigation or assurance for precluding that fault. Examples of this are shown in MG-13.
(iii) Mitigating features which include “shake and bake” testing on each LRU, dissimilar design, and architecture considerations such as simplex back-up systems.
(2) All radio and navigation systems and equipment should be installed and operated in a manner that will not have a detrimental effect on the proper functioning of any electronic equipment or system required by the FAR. It should be noted that §§ 29.1301 (reference paragraph AC 29.1301) and 29.1309(b) through (d) (reference paragraph AC 29.1309) apply to all installed equipment and systems and § 29.1309(a) Page F – 106
applies to all systems and equipment required by Parts 21 through 49. As an example of showing compliance with this section, consider a high frequency radio (HF) system installation. The first thing to determine is that the installation and operation of the HF system cannot create a hazard. Consideration may be necessary in hazardous situations such as precipitation on the antenna. Next, it should be determined that the operation of the HF does not cause interference to a system whose functioning is required by the FAR. An example of unacceptable interference would be if operating the HF transmitter caused one of the navigation radios to malfunction.
(3) Finally, it should be determined that other systems do not interfere with the HF system. Additional guidance on the testing of avionics equipment and installation is contained in paragraph AC 29 MG 1.
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AC 29-2C, Chg 3 9/30/2008 AC 29.1433. § 29.1433 VACUUM SYSTEMS.
a. Explanation. Vacuum systems have been utilized on some rotorcraft to provide an energy source for the flight instruments. This specific rule addresses the potential hazards which are peculiar to vacuum system installations. The possible fire hazards presented by these systems are of particular concern.
b. Procedure. The following items should be specifically addressed when evaluating a vacuum system installation:
(1) Pressure and Temperature Protection. The high- pressure outlet of the vacuum pump should have a means to automatically relieve the pressure if it becomes excessively high or the air temperature becomes excessively hot.
(2) Fire Hazard Protection. The components of the vacuum system that are mounted in a designated fire zone should be fire resistant. This includes engine or transmission driven pumps if they are in a fire zone. The discharge side of the pump may emit flammable fluids. This discharge side of the pump, along with its associated lines and fittings, should meet the criteria in paragraph AC 29.1183.
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AC 29.1435. § 29.1435 HYDRAULIC SYSTEM S.
a. Reference Regulations. The following sections of Part 29 are either incorporated in the provisions of § 29.1435 or are otherwise applicable to hydraulic system design:
(1) Section 29.695. Paragraph AC 29.695 covers power boost and power operated control systems.
(2) Section 29.861. Paragraph AC 29.861 covers fire protection of structure, controls, and other parts.
(3) Section 29.863. Paragraph AC 29.863 covers flammable fluid fire protection.
(4) Section 29.1183. Paragraph AC 29.1183 covers lines, fittings, and components.
(5) Section 29.1185. Paragraph AC 29.1185 covers flammable fluids.
(6) Section 29.1189. Paragraph AC 29.1189 covers shutoff means.
(7) Section 29.1309. Paragraph AC 29.1309 covers the requirements for functioning and reliability, and prevention of hazards if malfunctions or failures occur.
(8) Section 29.1322. Paragraph AC 29.1322 covers warning, caution, and advisory lights.
b. System Design. It is assumed that the hydraulic system is to be utilized to operate the primary control system of the rotorcraft and the rotorcraft cannot be safely operated without the hydraulic system.
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(1) Section 29.1309, paragraphs (a) and (b), provides for functioning reliably under any foreseeable operating condition and prevention of hazards after any malfunction or failure.
(2) The substantiating data should include a failure analysis that considers every possible system component failure, such as (but not limited to) ruptured lines, pump failure, regulator failure, ruptured seals, clogged filters, broken pilot valve connections, etc.
(3) The requirements of § 29.1309(a) and (b) are met by dual independent hydraulic systems from the reservoir, hydraulic pump, regulator, connecting tubing, and hoses through the actuators. There must be no commonality in the fluid- containing components. A break in one system should not result in fluid loss in the remaining system.
(4) The pumps should be separated as far as practicable; i.e., on opposite sides of the rotor drive transmission, on separate engines, or one pump on an engine and the other on the rotor drive transmission. The tubing and hoses should also be routed with as much physical separation as practicable. The purpose of this separation is to prevent total loss of the hydraulic systems in the event of a malfunction such as fire, or rotor burst wherein one projectile could disable both systems.
(5) Dual actuators must be designed to assure that any single failure, such as a cracked housing, broken interconnecting input, or output link, does not result in loss of total hydraulic system function.
(6) If the assumption under (b) above does not apply and the pilot can control the rotorcraft without undue fatigue after loss of the hydraulic system, then a single hydraulic system is acceptable.
(7) The pressure-indicating system required by § 29.1435, paragraph (a)(3), can be satisfied with a dial, vertical scale, or digital indicator. The indicator should enable the crew to detect pressure trends. Paragraph AC 29.1322 concerns § 29.1322 regarding proper colors for annunciators if they are used to supplement the indicating system.
(8) An analysis or a combination of analysis and tests must be included in the substantiating data file to show compliance with paragraphs (a)(1), (a)(2), and (a)(4) of § 29.1435.
(9) Extra caution should be exercised to assure that control input forces at the mechanical connection to the actuator pilot valves do not exceed their intended value.
Consideration should be given to the most adverse tolerance buildup in parts fabrication and control system rigging.
Page F – 116